Method and apparatus for reducing axial compressor blade tip flow

ABSTRACT

A turbine machine of the type having a high pressure compressor positioned in a casing is provided with a plurality of rotating compressor blades with at least one air channel formed in the blades for air flow communication from the base to the tip for extracting and pressurizing air from an inner area proximate the disk and conveying the extracted and pressurized air into an area between the blade tip and casing for blocking air flow across the tips of the blades.

TECHNICAL FIELD AND BACKGROUND OF THE INVENTION

This invention relates to a method and apparatus for reducing axialcompressor tip flow in airfoils, such as blades and vanes. Blade tipflow in the compressor area of a turbine engine results in loss ofcompressor efficiency and stall margin. In addition, flow recirculationin seal cavities along the inner flow path between the vanes and bladesalso degrades compressor performance. One prior art solution forreducing tip flow is to reduce the blade tip clearance. This is done bya variety of means, including control of the casing and vane interfaceusing mechanical and/or thermal methods. These methods can cause tiprubbing, excess wear and loss of engine efficiency.

BRIEF DESCRIPTION OF THE INVENTION

According to one aspect of the invention, a method of reducing air flowbetween a tip of a turbine airfoil rotating in a closely-spaced apartcasing is provided, and comprises the step of providing in the airfoil aradially-extending channel having an inlet opening proximate a base ofthe airfoil and an exit opening on the airfoil tip. Air is extracted andpressurized from a region proximate the base of the airfoil, introducedinto the channel and conveyed through the channel to the airfoil tip.The air exits the channel through the exit openings in the airfoil tipinto an area between the airfoil tip and casing under sufficientpressure to resist axial air flow from a pressure side to a suction sideof the airfoil.

Another aspect of the invention provides a method of reducing air flowbetween a tip of a turbine airfoil rotating in a closely-spaced apartcasing, comprising the steps of providing a first radially-extendingchannel having an inlet opening proximate a base of the airfoil on aleading edge side thereof, and an exit opening on the airfoil tip, andproviding a second radially-extending channel having an inlet openingproximate the base of the airfoil on a trailing edge side thereof, andan exit opening on the airfoil tip. The air is extracted from a regionproximate the base of the airfoil into the channel and pumped throughthe channel to the airfoil tip. The air exits the channel through theexit openings in the airfoil tip into an area between the airfoil tipand casing under sufficient pressure to resist axial air flow from apressure side to a suction side of the airfoil.

In another aspect of the invention a turbine machine compressor airfoilis provided, comprising a airfoil base, a airfoil tip, and an air flowchannel extending radially from an air inlet opening in the airfoilproximate the airfoil base to an exit opening in the airfoil tip forproviding a air blockage against an axial flow of air from the pressureside of the airfoil to the suction side of the airfoil to thereby reducecompressor airfoil tip flow.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is described below in conjunction with the followingdrawings, in which:

FIG. 1 is a fragmentary cross-section of an axial flow compressorsection of a turbine engine illustrating one embodiment of theinvention;

FIG. 2 is an enlarged fragmentary cross-section of a portion of thecompressor shown in FIG. 1;

FIG. 3 is a fragmentary cross-section of the compressor section of aturbine engine illustrating another embodiment of the invention;

FIG. 4 is an enlarged fragmentary cross-section of a portion of thecompressor shown in FIG. 3;

FIG. 5 is a fragmentary cross-section of the compressor section of aturbine engine illustrating yet another embodiment of the invention; and

FIG. 6 is an enlarged fragmentary cross-section of a portion of thecompressor shown in FIG. 5.

DESCRIPTION OF THE PREFERRED EMBODIMENT AND BEST MODE

Referring now specifically to the drawings, a partial section of theaxial compressor section of a turbine engine T1 illustrating a methodand apparatus for controlling axial compressor blade tip flow accordingto the present invention is illustrated in FIG. 1. The turbine engine“T1” includes compressor blades 10-14 and intermediately-positionedstator vanes 15-19 in a casing C1. The compressor blades 10-14 includerespective leading edges 10A-14A.

As is shown in FIG. 2, blade 10 is shown in enlarged detail for clarity,and is also exemplary of blades 11-14. Air is extracted and pressurizedfrom the area of the leading edge side 10A of the blade 10 through holes10B in a disk 20. The holes 10B communicate with a channel 10C thatextends radially outwardly through the blade 10 to the tip where itexits through holes 10D. Note that the channel 10C may branch out beforeexiting the tip of the blade 10. The size of the channel 10C and thelocation and number of the branches is determined empirically based onblade size, shape and volume, and engine performance, rating, tipclearance and similar factors. Note in the drawings that the tipclearance is sufficiently small in relation to the scale of the drawingsthat actual representation of the tip clearance cannot be shown.

Referring now to FIG. 3, a turbine engine “T2” includes compressorblades 30-34 and intermediately-positioned stator vanes 35-39 in acasing C2. The compressor blades 30-34 include respective trailing edges30A-34A.

In FIG. 4, blade 31 is shown in enlarged detail for clarity, and isexemplary of blades 30 and 32-34. Air is extracted from the area of thetrailing edge side 31A of the blade 31 through holes 31B in the disk rim40. The holes 31B communicate with a channel 31C that extends radiallyoutwardly through the blade 31 to the tip where it preferably branchesbefore exiting through holes 31D.

Referring now to FIG. 5, a turbine engine “T3” includes compressorblades 50-54 and intermediately-positioned stator vanes 55-59 in acasing C3. The compressor blades 50-54 include respective leading edges50A-54A and respective trailing edges 50B-54B.

FIG. 6 illustrates a blade 52 that is shown in enlarged detail forclarity, and is exemplary of blades 51 and 52-54. Air is extracted fromboth the areas of the leading edge side 52A and trailing edge side 52Bof blade 52 through holes 52C and 52D in the disk rim 60. The holes 52Cand 52D communicate with channels 52E and 52F, respectively, that extendradially outwardly through the blade 52 to the tip, where theypreferably branch before exiting through holes 52G.

In each of the embodiments described above, the air discharged at theblade tip reduces or prevents blade tip flow by aerodynamically blockingair flow in the region of the tip clearance between the blade tip andthe casing. Air from the inner flow path is brought to the tipclearance, as described above, to form this air block. The pressure ofthe extracted air increases due to the compressor rotor pumping and,when exiting the blade at the tip, resists air flow across the blade tipfrom the pressure side to the suction side.

The methods described above can be applied to both low pressurecompressors (boosters) and high pressure compressors. There is nochargeable flow loss when these methods are utilized. Furthermore, byreducing air flow by aerodynamic air blockage rather than by a tightrunning clearance between the blade tips and the casing, a largerassembly clearance between the blade tips and the casing can beestablished and maintained. Blade tip rubs are thus reduced, as isrecirculation in the inner flow path between the vane and the blade. Theextracted air is continuously pumped from the inner flow path to theblade tip, thus providing a continuous air blockage to the blade tip atall times during engine operation.

The methods described in this application also have application in blisk(blade integrated disk), skewed or circumferential dovetailed blades.

A method and apparatus for controlling axial compressor blade tip flowis described above. Various details of the invention may be changedwithout departing from its scope. Furthermore, the foregoing descriptionof the preferred embodiment of the invention and the best mode forpracticing the invention are provided for the purpose of illustrationonly and not for the purpose of limitation—the invention being definedby the claims.

1. A turbine machine of the type having a high pressure compressorpositioned in a casing, a plurality of rotating compressor blades havingrespective blade tips, respective blade bases affixed to a central disk,and a plurality of stationary vanes positioned between respective onesof the blades, comprising at least one air channel formed in respectiveones of the blades for air flow communication from the base to the tipfor extracting and pressurizing air from an inner area proximate thedisk and conveying the extracted and pressurized air through the channelinto an area between the blade tip and casing for blocking air flowacross a tip of the blades.
 2. A turbine machine according to claim 1,wherein the channel exits the blade tip through a respective pluralityof exit holes.
 3. A turbine machine according to claim 1, wherein thechannel includes an air inlet opening in on leading edge side of theblade.
 4. A turbine machine according to claim 1, wherein the channelincludes an air inlet opening on a trailing edge side of the blade.
 5. Aturbine machine according to claim 1, wherein the channel includes anair inlet opening in on a leading edge side and an air inlet opening ona trailing edge side of the blade.
 6. A turbine machine according toclaim 1, and including a first channel having an air inlet opening on aleading edge side of the blade and a second channel having an air inletopening on a trailing edge side of the blade.
 7. A turbine machineaccording to claims 3, wherein the channel includes a plurality of exitholes in the blade tip.
 8. A method of reducing air flow between a tipof a turbine airfoil rotating in a closely-spaced apart casing,comprising the steps of: (a) providing in the airfoil aradially-extending channel having an inlet opening proximate a base ofthe airfoil and an exit opening on the airfoil tip; (b) extracting andpressurizing air from a region proximate the base of the airfoil intothe channel; (c) conveying the air through the channel to the airfoiltip; and (d) discharging the air through the exit openings in theairfoil tip into an area between the airfoil tip and casing undersufficient pressure to resist axial air flow across the tip of theairfoil.
 9. A method according to claim 8, wherein the step of providinga radially-extending channel having an inlet opening proximate a base ofthe airfoil and an exit opening on the airfoil tip includes the step offorming the inlet opening on a leading edge side of the airfoil.
 10. Amethod according to claim 8, wherein the step of providing aradially-extending channel having an inlet opening proximate a base ofthe airfoil and an exit opening on the airfoil tip includes the step offorming the inlet opening on a trailing edge side of the blade.
 11. Amethod according to claim 8, wherein the step of providing aradially-extending channel having an inlet opening proximate a base ofthe airfoil and an exit opening on the airfoil tip includes the steps offorming an air inlet opening on a leading edge side of the airfoil, andan air inlet opening on the trailing edge side of the airfoil.
 12. Amethod according to claim 8, including the steps of providing: (a) afirst radially-extending channel having an inlet opening proximate thebase of the airfoil on the leading edge side thereof and an exit openingat the airfoil tip; and (b) a second radially-extending channel havingan inlet opening proximate the base of the airfoil on the trailing edgeside thereof and an exit opening at the airfoil tip.
 13. A method ofreducing air flow between a tip of a turbine airfoil rotating in aclosely-spaced apart casing, comprising the steps of: (a) providing afirst radially-extending channel having an inlet opening proximate abase of the airfoil on a leading edge side thereof, and an exit openingon the airfoil tip; (b) providing a second radially-extending channelhaving an inlet opening proximate the base of the tip on a trailing edgeside thereof, and an exit opening on the airfoil tip; (c) extractinghigh pressure air from a region proximate the base of the airfoil intothe channels: (d) conveying the air through the channels to the airfoiltip; and (e) exiting the air through the exit openings in the airfoiltip into an area between the airfoil tip and casing under sufficientpressure to resist axial air flow across the tip of the airfoil.
 14. Aturbine machine compressor blade, comprising: (a) a blade base; (b) ablade tip; (c) an air flow channel extending radially from an air inletopening in the blade proximate the blade base to an exit opening in theblade tip.
 15. A turbine machine compressor blade according to claim 14,wherein the air inlet opening is formed on a leading edge side of theblade.
 16. A turbine machine compressor blade according to claim 14,wherein the air inlet opening is formed on a trailing edge side of theblade.
 17. A turbine machine compressor blade according to claim 14,wherein the channel includes an air inlet opening on a leading edge sideof the blade and an air inlet opening on a trailing edge side of theblade.
 18. A turbine machine compressor blade according to claim 14, andincluding a first channel having an air inlet opening on a leading edgeside of the blade and a second channel having an air inlet opening on atrailing edge side of the blade.